Gas turbine engines operate at high temperatures in order to increase their efficiency. Various advancements have been employed to enable the components (e.g., airfoils) of such engines to operate for longer periods of time at such high temperature. Airfoils employed in modern, high efficiency power generation combustion turbine engines rely on high quality materials such as single crystal alloys and precise control of the part's internal and external dimensions. In addition to the use of high temperature resistant super alloys, various airfoils have been designed to include internal cooling systems. One such internal cooling system is the use of cooling passages located inside and near the surface of the airfoil.
A number of techniques have been employed to provide such turbine airfoils with near surface cooling passages. For example, U.S. Pat. No. 6,638,639 discloses high efficiency, thin-walled turbine components such as turbine blade airfoils comprising a superalloy substrate with cooling channels covered by a thin super alloy skin. The thin skin is bonded to the inner spar structure of a turbine blade airfoil. U.S. Pat. No. 6,321,449 discloses a method of forming an internal channel within an article, such as a cooling channel in an air-cooled blade, vane, shroud, combustor or duct of a gas turbine engine. The method generally entails forming a substrate to have a groove recessed in its surface. A sacrificial material is deposited in the groove to form a filler that can be preferentially removed from the groove. A permanent layer is deposited on the surface of the substrate and over the filler, after which the filler is removed from the groove to yield the desired channel in the substrate beneath the permanent layer.
The present invention is an improvement over known techniques for providing turbine airfoils with near surface cooling capability.